Geared turbofan engine with a high ratio of thrust to turbine volume

ABSTRACT

A gas turbine engine turbine has a high pressure turbine configured to rotate with a high pressure compressor as a high pressure spool in a first direction about a central axis and a low pressure turbine configured to rotate with a low pressure compressor as a low pressure spool in the first direction about the central axis. A power density is greater than or equal to about 1.5 and less than or equal to about 5.5 lbf/cubic inches. A fan is connected to the low pressure spool via a speed changing mechanism and rotates in the first direction.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. application Ser. No.17/192,039 filed on Mar. 4, 2021, which is continuation of U.S.application Ser. No. 16/158,545 filed on Oct. 12, 2018, now U.S. Pat.No. 11,053,843 granted on Jul. 6, 2021, which is continuation of U.S.application Ser. No. 14/592,991 filed on Jan. 9, 2015, now U.S. Pat. No.10,138,809 granted on Nov. 27, 2018, which is a continuation-in-part ofU.S. application Ser. No. 13/445,095, filed Apr. 12, 2012, which claimsthe benefit of U.S. Provisional Ser. No. 61/619,133, which was filedApr. 2, 2012.

BACKGROUND OF THE INVENTION

This application relates to a geared turbofan gas turbine engine,wherein the low and high pressure spools rotate in the same directionrelative to each other.

Gas turbine engines are known, and typically include a fan deliveringair into a compressor section, and outwardly as bypass air to providepropulsion. The air in the compressor is delivered into a combustionsection where it is mixed with fuel and burned. Products of thiscombustion pass downstream over turbine rotors, driving them to rotate.Typically there are low and high pressure compressors, and low and highpressure turbines.

The high pressure turbine typically drives the high pressure compressoras a high spool, and the low pressure turbine drives the low pressurecompressor and the fan. Historically, the fan and low pressurecompressor were driven at a common speed.

More recently, a gear reduction has been provided on the low pressurespool such that the fan and low pressure compressor can rotate atdifferent speeds. It desirable to have more efficient engines that havemore compact turbines to limit efficiency loses.

SUMMARY

In a featured embodiment, a gas turbine engine turbine comprises a highpressure turbine configured to rotate with a high pressure compressor asa high pressure spool in a first direction about a central axis. A lowpressure turbine is configured to rotate in the first direction aboutthe central axis. A fan is connected to the low pressure turbine via agear reduction and will rotate in the first direction. The engine isconfigured to have a ratio of a thrust provided by the engine, to avolume of a turbine section including both the high pressure turbine andthe low pressure turbine, that is greater than or equal to about 1.5 andless than or equal to about 5.5 lbf/in³. The thrust is sea leveltake-off, flat-rated static thrust.

In another embodiment according to the previous embodiment, guide vanesare positioned upstream of a first stage in the low pressure turbine todirect gases downstream of the high pressure turbine as the gasesapproach the low pressure turbine.

In another embodiment according to any of the previous embodiments, amid-turbine frame supports the high pressure turbine.

In another embodiment according to any of the previous embodiments, theguide vanes are positioned intermediate the mid-turbine frame and thelow pressure turbine.

In another embodiment according to any of the previous embodiments,there is an intermediate section, and the intermediate turbine sectiondrives a compressor rotor.

In another embodiment according to any of the previous embodiments, thegear reduction is positioned intermediate the fan and a compressor rotordriven by the low pressure turbine.

In another embodiment according to any of the previous embodiments, thegear reduction is positioned intermediate the low pressure turbine and acompressor rotor driven by the low pressure turbine.

These and other features may be best understood from the followingdrawings and specification.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically shows a gas turbine engine.

FIG. 2 schematically shows rotational features of one type of such anengine.

FIG. 3 is a detail of the turbine section volume.

FIG. 4 shows another embodiment.

FIG. 5 shows yet another embodiment.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude, for example, three-spools, an augmentor section, or a differentarrangement of sections, among other systems or features. The fansection 22 drives air along a bypass flowpath B while the compressorsection 24 drives air along a core flowpath C for compression andcommunication into the combustor section 26 then expansion through theturbine section 28. The bypass flowpath is defined between a fan housingoutward of fan section 22 and a core housing. The core flowpath iswithin the core housing. Although depicted as a turbofan gas turbineengine in the disclosed non-limiting embodiment, it should be understoodthat the concepts described herein are not limited to use with turbofansas the teachings may be applied to other types of turbine engines. Forpurposes of this application, the terms “low” and “high” as applied tospeed or pressure are relative terms. The “high” speed and pressurewould be higher than that associated with the “low” spools, compressorsor turbines, however, the “low” speed and/or pressure may actually be“high.”

The engine 20 generally includes a low speed spool 30 and a high speedspool 32 mounted for rotation about an engine central longitudinal axisA relative to an engine static structure 36 via several bearing systems38. It should be understood that various bearing systems 38 at variouslocations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through ageared architecture 48 to drive the fan 42 at a lower speed than the lowspeed spool 30. The high speed spool 32 includes an outer shaft 50 thatinterconnects a high pressure compressor 52 and high pressure turbine54. The terms “high” and “low” in relation to both the speed andpressure of the components are relative to each other, and not to anabsolute value. A combustor 56 is arranged between the high pressurecompressor 52 and the high pressure turbine 54. A mid-turbine frame 57of the engine static structure 36 is arranged generally between the highpressure turbine 54 and the low pressure turbine 46. The mid-turbineframe 57 further supports bearing systems 38 in the turbine section 28.The inner shaft 40 and the outer shaft 50 are concentric and rotate viabearing systems 38 about the engine central longitudinal axis A which iscollinear with their longitudinal axes.

The core airflow C is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path and act as inlet stator vanes to turn theflow to properly feed the first blades of the low pressure turbine. Theturbines 46, 54 rotationally drive the respective low speed spool 30 andhigh speed spool 32 in response to the expansion.

The engine 20 has bypass airflow B, and in one example is a high-bypassgeared aircraft engine. The bypass ratio may be defined as the amount ofair delivered into the bypass duct divided by the amount delivered intothe core flow. In a further example, the engine 20 bypass ratio isgreater than about six (6), with an example embodiment being greaterthan ten (10), the geared architecture 48 is an epicyclic gear train,such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3 and the low pressure turbine46 has a pressure ratio that is greater than about 5. In one disclosedembodiment, the engine 20 bypass ratio is greater than about ten (10:1),the fan diameter is significantly larger than that of the low pressurecompressor 44, and the low pressure turbine 46 and the low pressureturbine has a pressure ratio that is greater than about 5:1. Lowpressure turbine 46 pressure ratio is the total pressure measured priorto inlet of low pressure turbine 46 as related to the pressure at theoutlet of the low pressure turbine 46 prior to an exhaust nozzle. Thegeared architecture 48 may be a planet gear arrangement such that thefan will rotate in the same direction as the low spool. It should beunderstood, however, that the above parameters are only exemplary of oneembodiment of a geared architecture engine and that the presentinvention is applicable to other gas turbine engines including directdrive turbofans.

A greatest amount of thrust is provided by the bypass flow B due to thehigh bypass ratio. The fan section 22 of the engine 20 is designed for aparticular flight condition—typically cruise at about 0.8 Mach and about35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, with theengine at its best fuel consumption—also known as “bucket cruise ThrustSpecific Fuel Consumption (‘TSFC’)”—is the industry standard parameterof lbm of fuel being burned per hour divided by lbf of thrust the engineproduces at that minimum point. “Low fan pressure ratio” is the pressureratio across the fan blade alone, before the Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tram degR)/518.7){circumflex over ( )}0.5]. The “Low corrected fan tip speed” asdisclosed herein according to one non-limiting embodiment is less thanabout 1150 ft/second at the same cruise point.

FIG. 2 shows detail of an engine 120, which may generally have thefeatures of engine 20 of FIG. 1 . A fan 122 is positioned upstream of alow pressure compressor 124, which is upstream of a high pressurecompressor 126. A combustor 128 is positioned downstream of the highpressure compressor 126. A mid-turbine frame 142 may be positioned at adownstream end of the high pressure turbine 130, and supports a bearing138, shown schematically, to support the aft end of the high pressureturbine 130, and a high pressure spool 132. A low pressure turbine 134is positioned downstream of a mid-turbine frame 142. A low spool 136,driven by the low pressure turbine 134, drives the low pressurecompressor 124. The speed change mechanism 48 causes the fan 122 torotate at a different speed than the low pressure compressor 134. Inembodiments of this invention, the speed input to output ratio for thespeed change mechanism is above or equal to 2.3:1, and up to less thanor equal to 13:1. The gear also causes fan 122 to rotate in the samedirection relative to the low pressure compressor 124. As mentionedabove, a planet gear arrangement may be utilized to cause the fan 122 torotate in the same direction (“-”) relative to the low pressurecompressor 124. In this embodiment the fan generally has less than 26blades, and the low pressure turbine has at least three stages, and upto six stages. The high pressure turbine generally has one or two stagesas shown.

In this particular embodiment, the low pressure compressor 124 and thelow pressure turbine 134 rotate in one direction (“-”) and the highpressure turbine 130, the high pressure compressor 126, rotate in thesame direction (“-”) as does fan 122.

A strut 140 is shown between the low pressure compressor 124 and thehigh pressure compressor 126. The strut 140 spans the gas path, and hasan airfoil shape, or at least a streamline shape. The combination of ablade at the exit of the low pressure compressor 124, the strut 140, anda variable vane, and then the first blade of the high pressurecompressor 126 is generally encompassed within the structure illustratedas the strut 140.

Since the compressor sections 124 and 126 rotate in the same direction,the several airfoils illustrated as the element 140 are required to doless turning of the air flow.

As will be explained below, since the turbine section is provided with ahighly cambered vane, there is less turning required between the twoturbine sections. Since the compressor is forcing flow with an adversepressure gradient, and whereas the turbine has a favorable pressuregradient, this overall engine architecture is benefited by theillustrated combination.

Highly cambered inlet guide vanes 143 are positioned in a locationintermediate the mid-turbine frame 142 and the most upstream rotor inthe low pressure turbine 134. The vanes 143 must properly direct theproducts of combustion downstream of the high pressure turbine 130 asthey approach the first rotor of the low pressure turbine 134. It isdesirable for reducing the overall size of the low pressure turbine thatthe flow be properly directed when it initially encounters the firststage of the low pressure turbine section.

The above features achieve a more compact turbine section volumerelative to the prior art, including both the high and low pressureturbines. A range of materials can be selected. As one example, byvarying the materials for forming the low pressure turbine, the volumecan be reduced through the use of more expensive and more exoticengineered materials, or alternatively, lower priced materials can beutilized. In three exemplary embodiments the first rotating blade of theLow Pressure Turbine can be a directionally solidified cast blade, asingle crystal cast blade or a hollow, internally cooled blade. Allthree embodiments will change the turbine volume to be dramaticallysmaller than the prior art by increasing low pressure turbine speed. Inaddition, high efficiency blade cooling may be utilized to furtherresult in a more compact turbine section.

Due to the compact turbine section, a power density, which may bedefined as thrust in pounds force produced divided by the volume of theentire turbine section, may be optimized. The volume of the turbinesection may be defined by an inlet of a first turbine vane in the highpressure turbine to the exit of the last rotating airfoil in the lowpressure turbine, and may be expressed in cubic inches. The staticthrust at the engine's flat rated Sea Level Takeoff condition divided bya turbine section volume is defined as power density. The sea leveltake-off flat-rated static thrust may be defined in lbs force, while thevolume may be the volume from the annular inlet of the first turbinevane in the high pressure turbine to the annular exit of the downstreamend of the last rotor section in the low pressure turbine. The maximumthrust may be Sea Level Takeoff Thrust “SLTO thrust” which is commonlydefined as the flat-rated static thrust produced by the turbofan atsea-level.

The volume V of the turbine section may be best understood from FIG. 3 .As shown, the frame 142 and vane 143 are intermediate the high pressureturbine section 130, and the low pressure turbine section 134. Thevolume V is illustrated by dashed line, and extends from an innerperiphery I to an outer periphery O. The inner periphery is somewhatdefined by the flowpath of the rotors, but also by the inner platformflow paths of vanes. The outer periphery is defined by the stator vanesand outer air seal structures along the flowpath. The volume extendsfrom a most upstream end of the vane 400 at the beginning of the highpressure turbine 130, typically its leading edge, and to the mostdownstream edge 401 of the last rotating airfoil in the low pressureturbine section 134. Typically this will be the trailing edge of thatairfoil.

The power density in the disclosed gas turbine engine is much higherthan in the prior art. Eight exemplary engines are shown below whichincorporate turbine sections and overall engine drive systems andarchitectures as set forth in this application, and can be found inTable I as follows:

TABLE 1 Thrust Turbine section Thrust/turbine SLTO volume from sectionvolume Engine (lbf) the Inlet (lbf/in³) 1 17,000 3,859 4.41 2 23,3005,330 4.37 3 29,500 6,745 4.37 4 33,000 6,745 4.84 5 96,500 31,086 3.106 96,500 62,172 1.55 7 96,500 46,629 2.07 8 37,098 6,745 5.50

Thus, in embodiments, the power density would be greater than or equalto about 1.5 lbf/in³. More narrowly, the power density would be greaterthan or equal to about 2.0 lbf/in³.

Even more narrowly, the power density would be greater than or equal toabout 3.0 lbf/in³.

More narrowly, the power density is greater than or equal to about 4.0lbf/in³. More narrowly, the power density is greater than or equal toabout 4.5 lbf/in³. Even more narrowly, the power density is greater thanor equal to about 4.75 lbf/in³. Even more narrowly, the power density isgreater than or equal to about 5.0 lbf/in³.

Also, in embodiments, the power density is less than or equal to about5.5 lbf/in³.

While certain prior engines have had power densities greater than 1.5,and even greater than 3.2, such engines have been direct drive enginesand not associated with a gear reduction. In particular, the powerdensity of an engine known as PW4090 was about 1.92 lbf/in³, while thepower density of an engine known as V2500 had a power density of 3.27lbf/in³.

Engines made with the disclosed architecture, and including turbinesections as set forth in this application, and with modifications comingfrom the scope of the claims in this application, thus provide very highefficient operation, and increased fuel efficiency and lightweightrelative to their trust capability.

FIG. 4 shows an embodiment 200, wherein there is a fan drive turbine 208driving a shaft 206 to in turn drive a fan rotor 202. A gear reduction204 may be positioned between the fan drive turbine 208 and the fanrotor 202. This gear reduction 204 may be structured and operate likethe gear reduction disclosed above. A compressor rotor 210 is driven byan intermediate pressure turbine 212, and a second stage compressorrotor 214 is driven by a turbine rotor 216. A combustion section 218 ispositioned intermediate the compressor rotor 214 and the turbine section216.

FIG. 5 shows yet another embodiment 300 wherein a fan rotor 302 and afirst stage compressor 304 rotate at a common speed. The gear reduction306 (which may be structured as disclosed above) is intermediate thecompressor rotor 304 and a shaft 308 which is driven by a low pressureturbine section.

The FIG. 4 or 5 engines may be utilized with the density featuresdisclosed above.

Although an embodiment of this invention has been disclosed, a person ofordinary skill in this art would recognize that certain modificationswould come within the scope of this application. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this invention.

1. A gas turbine engine turbine comprising: a high pressure turbineconfigured to rotate with a high pressure compressor about a centralaxis; a low pressure turbine configured to rotate with a low pressurecompressor about said central axis; a propulsor having a plurality ofpropulsor blades and connected to the low pressure turbine via a gearreduction, wherein the gas turbine engine is configured to have a powerratio of a thrust provided by said engine, to a volume of a turbinesection including both said high pressure turbine and said low pressureturbine that is greater than or equal to 3.0 and less than or equal to5.5 lbf/in³; said thrust is sea level take-off, flat-rated staticthrust; wherein said low pressure turbine having a pressure ratiogreater than 5.0, defined by a total pressure measured prior to an inletof said low pressure turbine as related to a pressure at an outlet ofsaid low pressure turbine prior to an exhaust nozzle; and a mid-turbineframe positioned intermediate said high pressure turbine and said lowpressure turbine, said mid-turbine frame including a bearing, whereinsaid bearing support said high pressure turbine, and said mid-turbineframe including airfoils in a core airflow path that provide inletstator vanes to turn a flow to feed a first blade row of said lowpressure turbine.
 2. The gas turbine engine as set forth in claim 1,wherein said low pressure turbine has between three and six stages, andsaid high pressure turbine has one or two stages.
 3. The gas turbineengine as set forth in claim 2, wherein a gear ratio of said gearreduction being greater than 2.3.
 4. The gas turbine engine as set forthin claim 3, wherein the low pressure turbine, the high pressure turbineand the propulsor all rotate in a common direction.
 5. The gas turbineengine as set forth in claim 4, wherein the gear reduction is positionedintermediate the propulsor and the low pressure compressor.
 6. The gasturbine engine as set forth in claim 1, wherein said power ratio isgreater than or equal to 4.0 lbf/in³.
 7. The gas turbine engine as setforth in claim 6, wherein said power ratio is greater than or equal to4.5 lbf/in³.
 8. The gas turbine engine as set forth in claim 6, whereinsaid plurality of propulsor blades being less than twenty six blades. 9.The gas turbine engine as set forth in claim 8, wherein turbine bladesin a first row of said low pressure turbine being at least one ofdirectionally solidified cast blades, a single crystal cast blade, or ahollow internally cooled blade.
 10. The gas turbine engine as set forthin claim 1, wherein there are turbine blades in a first blade row ofsaid low pressure turbine that are at least one of directionallysolidified cast blades, a single crystal cast blade, or a hollowinternally cooled blade.
 11. The gas turbine engine as set forth inclaim 10, wherein said low pressure turbine has between three and sixstages, and said high pressure turbine has one or two stages.
 12. Thegas turbine engine as set forth in claim 11, wherein said propulsor is afan received within a fan housing.
 13. The gas turbine engine as setforth in claim 1, wherein said propulsor is a fan received within a fanhousing.
 14. A gas turbine engine turbine comprising: a high pressureturbine configured to rotate with a high pressure compressor about acentral axis; a low pressure turbine configured to rotate with a lowpressure compressor about said central axis; a propulsor having aplurality of propulsor blades and connected to the low pressure turbinevia a gear reduction, wherein the gas turbine engine is configured tohave a power ratio of a thrust provided by said engine, to a volume of aturbine section including both said high pressure turbine and said lowpressure turbine that is greater than or equal to 1.5 and less than orequal to 5.5 lbf/in³; said thrust is sea level take-off, flat-ratedstatic thrust; and said low pressure turbine has between three and sixstages, and said high pressure turbine has one or two stages.
 15. Thegas turbine engine as set forth in claim 14, wherein there are turbineblades in a first blade row of said low pressure turbine that are atleast one of directionally solidified cast blades, a single crystal castblade, or a hollow internally cooled blade
 16. The gas turbine engine asset forth in claim 15, wherein a gear ratio of said gear reduction beinggreater than 2.3.
 17. The gas turbine engine as set forth in claim 16,wherein said power ratio is greater than or equal to 4.0 lbf/in³. 18.The gas turbine engine as set forth in 16, wherein said low pressureturbine having a pressure ratio greater than 5.0, defined by a totalpressure measured prior to an inlet of said low pressure turbine asrelated to a pressure at an outlet of said low pressure turbine prior toan exhaust nozzle.
 19. The gas turbine engine as set forth in claim 18,a mid-turbine frame positioned intermediate said high pressure turbineand said low pressure turbine, said mid-turbine frame including abearing, wherein said bearing support said high pressure turbine, andsaid mid-turbine frame including airfoils in a core airflow path thatprovide inlet stator vanes to turn a flow to feed a first blade row ofsaid low pressure turbine
 20. The gas turbine engine as set forth inclaim 19, wherein said plurality of propulsor blades being less thantwenty six blades.
 21. The gas turbine engine as set forth in claim 15,a mid-turbine frame positioned intermediate said high pressure turbineand said low pressure turbine, said mid-turbine frame including abearing, wherein said bearing support said high pressure turbine, andsaid mid-turbine frame including airfoils in a core airflow path thatprovide inlet stator vanes to turn a flow to feed a first blade row ofsaid low pressure turbine.
 22. The gas turbine engine as set forth inclaim 14, a mid-turbine frame positioned intermediate said high pressureturbine and said low pressure turbine, said mid-turbine frame includinga bearing, wherein said bearing support said high pressure turbine, andsaid mid-turbine frame including airfoils in a core airflow path thatprovide inlet stator vanes to turn a flow to feed a first blade row ofsaid low pressure turbine.
 23. The gas turbine engine as set forth inclaim 14, wherein said power ratio is greater than or equal to 4.0lbf/in³.
 24. The gas turbine engine as set forth in claim 23, whereinsaid power ratio is greater than or equal to 4.5 lbf/in³.
 25. A gasturbine engine turbine comprising: a high pressure turbine configured torotate with a high pressure compressor about a central axis; a lowpressure turbine configured to rotate with a low pressure compressorabout said central axis; a propulsor having a plurality of propulsorblades and connected to the low pressure turbine via a gear reduction,wherein the engine is configured to have a power ratio of a thrustprovided by said engine, to a volume of a turbine section including bothsaid high pressure turbine and said low pressure turbine that is greaterthan or equal to 3.0 and less than or equal to 5.5 lbf/in³; said thrustis sea level take-off, flat-rated static thrust; wherein said lowpressure turbine having a pressure ratio greater than 5.0, defined by atotal pressure measured prior to an inlet of said low pressure turbineas related to a pressure at an outlet of said low pressure turbine priorto an exhaust nozzle; and said low pressure turbine has between threeand six stages, and said high pressure turbine has one or two stages.26. The gas turbine engine as set forth in claim 25, wherein said powerratio is greater than or equal to 4.0 lbf/in³.
 27. The gas turbineengine as set forth in claim 26, wherein said plurality of propulsorblades being less than twenty six blades.
 28. The gas turbine engine asset forth in claim 27, wherein there are turbine blades in a first bladerow of said low pressure turbine that are at least one of directionallysolidified cast blades, a single crystal cast blade, or a hollowinternally cooled blade.
 29. The gas turbine engine as set forth inclaim 25, wherein said plurality of propulsor blades being less thantwenty six blades.
 30. The gas turbine engine as set forth in claim 25,wherein there are turbine blades in a first blade row of said lowpressure turbine that are at least one of directionally solidified castblades, a single crystal cast blade, or a hollow internally cooledblade.